There is a growing trend in the aerospace industry to expand the use of composite materials for a diverse array of structural and dynamic applications because of the strength-to-weight advantage provided by composite materials. One particular application for the use of composite materials lies in the fabrication of composite articles such as structural panels, e.g., fuselage panels, for aircraft and helicopters that include one or more stiffening members for reacting loads experienced by the structural panel. Such structural panels generally comprise inner and outer composite skins, which are formed from composite materials such as fiberglass, graphite, or KEVLAR.RTM. (KEVLAR is a registered trademark of E. I. du Pont de Nemours & Co., Wilmington, Del. for an aromatic polyamide fiber of high tensile strength) embedded in a resinous matrix, e.g., epoxy, having a honeycomb core material interposed therebetween.
A structural panel typically includes one or more stiffening members affixed to the inner wall thereof for efficiently transmitting and/or reacting axial and/or bending loads to which the structural panel is subjected. For structural panels fabricated from sheet metal, the stiffening members are affixed to the inner wall by fasteners or welding. For structural panels fabricated from composite materials, composite stiffening members are typically affixed to the inner mold line (IML) composite skin of the composite structural panel by means of fasteners or bonding. While the use of fasteners is an acceptable technique for affixing composite stiffening members to the IML composite skin of a composite structural panel, this technique is extremely labor intensive, i.e., costly and time consuming, and the rivets utilized to effect affixation increase the overall weight attributable to the composite structural panel.
There are two techniques currently employed for bonding a composite stiffening member in combination with a composite structural panel: (1) the co-cured bonding method; or (2) the secondary bonding method. Both methods are disadvantageous in requiring costly non-recurring tooling and/or costly recurring manufacturing steps as discussed in further detail in the following paragraphs. FIG. 1A illustrates the co-cure bonding method for bonding a composite stiffening member SM in combination with a composite structural panel SP that includes an inner mold line (IML) composite skin S.sub.IML formed from composite plies and an outer mold line (OML) composite skin S.sub.OML formed from composite plies and which has a honeycomb core HC interposed between the OML and IML composite skins S.sub.OML, S.sub.IML.
A core mandrel CM is required to stabilize the composite plies P.sub.SM that comprise the stiffening member SM to prevent collapse or deformation of the plies P.sub.SM during the co-cure cycle as a result of the co-cure pressure exerted thereagainst. Typically, the core mandrel CM is formed from a lightweight, relatively rigid (relatively rigid being used herein in the sense that the core mandrel CM will accommodate, without collapse or deformation, the pressures experienced during the co-cure cycle while concomitantly having a minimal density) material, e.g., a rigid foam such as polyurethane foam. Recurring process steps are required to machine a block of the foam material to the net shape of the core mandrel CM. To ensure exact positioning of the core mandrel CM on the IML composite skin S.sub.IML is maintained during the co-cure cycle, a film adhesive FM is interposed between the core mandrel CM and the IML composite skin S.sub.IML. Recurring manufacturing steps are required to properly position the film adhesive FM on the IML composite skin S.sub.IML and to attach the core mandrel CM in aligned combination with the film adhesive FM. The attachment of core mandrel CM to the IML composite skin S.sub.IML by means of the film adhesive FM creates a slip-plane, i.e., a local weak joint or surface, along which the core mandrel CM may slide relative to the IML composite skin S.sub.IML. This creates the need for additional recurring tooling, i.e., a bonding fixture, during the co-cure cycle to ensure that no slippage of the core mandrel CM occurs due to the co-cure pressure exerted during the co-cure cycle.
FIG. 1B illustrates the secondary bonding method for bonding a precured composite stiffening member CSM in combination with the composite plies defining the composite structural panel SP. The precured composite stiffening member CSM is fabricated in prefabrication recurring manufacturing steps using non-recurring tooling. Non-recurring autoclave tooling, i.e., an autoclave molding assembly, is required to define the configuration of the composite stiffening member CSM to be prefabricated. Recurring manufacturing steps are required to lay up composite plies defining the composite stiffening member CSM in the autoclave tooling, and to cure the layed-up autoclave tooling in an autoclave to form the precured composite stiffening member CSM. Exact positioning of the precured composite stiffening member CSM on the IML composite skin S.sub.IML is maintained during the fabrication curing cycle wherein the OML and IML composite skins S.sub.OML, S.sub.IML are cured to form the finished composite structural panel SP by a combination of the inherent tackiness of the prepreg composite plies forming the IML composite skin S.sub.IML and a bonding fixture. While the secondary bonding method provides a marginal weight savings in the finished composite structural panel SP due to the elimination of the requirement for use of a core mandrel as a stabilizing member, the non-recurring tooling costs and recurring manufacturing steps required by this method, i.e., to prefabricate the composite stiffening member CSM, are more costly and/or time consuming than the non-recurring tooling costs and recurring process steps associated with the co-cure bonding method described hereinabove.
A need exists to provide a method for co-cure bonding a composite stiffening member in integral combination with a composite article that reduces the use of costly non-recurring tooling and/or reduces recurring manufacturing steps. The method should ensure the precise positioning of the composite stiffening member with respect to the honeycomb core of the composite article to be fabricated. The method should provide stabilization for the composite stiffening member during the co-cure cycle, i.e., preclude the collapse or deformation of the composite plies forming the composite stiffening member. Further, the method should provide stabilization of the honeycomb core during the co-cure cycle to preclude the collapse or deformation thereof.